The present invention generally relates to high pressure systems in rockets (e.g., rocket engines, oxidizer systems, propellant tanks) and, more particularly, to a vent system for such high pressure systems.
Multi-stage rockets have long been used to propel various types a spacecraft into outer space, including satellites. One known multi-stage rocket configuration has an upper stage whose forward end is attached to the spacecraft and whose aft end is interconnected with a lower stage by an interstage adapter. Prior to separation of the lower stage from the upper stage during flight and also prior to ignition of the upper stage engine, a flow of rocket fuel into the upper stage engine is initiated. This flow of rocket fuel is directed out of the upper stage engine through an upper stage rocket engine vent system. The upper stage rocket engine vent system directs this rocket fuel overboard via an appropriate conduit that passes through an exterior wall of the rocket (e.g., on the interstage adapter).
The lower stage is separated from the upper stage rocket engine at the appropriate time during flight, and the upper stage rocket engine is also ignited at the appropriate time. Further travel will then be affected by the thrust provided by the upper stage rocket engine. One prior art protocol that ultimately leads to the separation of the upper stage from the spacecraft entails shutting down the upper stage rocket engine a number of times. That is, the upper stage rocket engine is shut down for a predetermined period of time, is thereafter ignited and run for a predetermined period of time, and is thereafter shut down once again. Fuel is discharged from the upper stage rocket engine through the upper stage rocket engine vent system under very high pressure each time that the upper stage rocket engine is shut down. One prior art configuration for this upper stage rocket engine fuel vent system uses a pair of about 2 inch diameter cylindrical ducts that are interconnected with the upper stage rocket engine. Both of these ducts extend out from the rocket engine at least generally perpendicular to the primary or longitudinal axis of the upper stage, and thereafter are directed rearwardly (e.g., in a generally L-shaped configuration). The thrust that is generated out of each these cylindrical vent ducts is on the order of about 600 pounds in one prior art embodiment. Thrusts of this magnitude cause relatively significant movement of the ducts relative to the upper stage rocket engine and introduces significant structural stresses at the interconnection with the upper stage rocket engine. There is a strong potential for damage to the upper stage rocket engine because of the stresses.
The present invention generally relates to a system for reducing axial thrust for typically high pressure discharges. One particularly desirable application of the present invention is in a rocket engine vent system.
A first aspect of the present invention is embodied by a rocket that includes a spacecraft, as well a first or an upper stage that is interconnected with this spacecraft and that includes at least one rocket engine. A first or an upper stage rocket engine vent system is fluidly interconnected with this upper stage rocket engine (or any other rocket engine(s) associated with any other stage used by the rocket) and includes at least one first or upper stage rocket engine vent system conduit (hereafter a xe2x80x9cfirst rocket engine vent system conduitxe2x80x9d). A pair of vent apertures extend entirely through this first rocket engine vent system conduit. A first flow diverter that includes first and second vanes is disposed within this first rocket engine vent system conduit such that each vane projects at least generally toward its own vent aperture. Therefore, at least two separate and discrete fluid flows may be directed out of the first rocket engine vent system conduit.
Various refinements exist of the features noted in relation to the subject first aspect of the present invention. Further features may also be incorporated in the subject first aspect of the present invention as well. These refinements and additional features may exist individually or in any combination. The static pressure profile throughout the first rocket engine vent system conduit may be at least substantially unchanged (in relation to such a conduit that does not include the first flow diverter and vent apertures) by having the collective area of the two vent apertures (i.e., the sum of the surface area occupied by each of these vent apertures) be greater than the cross-sectional area of the first rocket engine vent system conduit (i.e., the area of the first rocket engine vent system conduit, taken perpendicular to a center reference axis about which the first rocket engine vent system conduit may be disposed). If a significant back pressure did develop in the first rocket engine vent system conduit, that is if the venting did not occur quickly enough such that this conduit (as well as the upper stage rocket engine) became overpressurized or such that the static pressure therein significantly increased, the upper stage rocket engine could be damaged. The pair of vent apertures are also preferably disposed in at least substantially opposing relation (e.g., 180 degrees apart), or stated another way on opposite sides of the first rocket vent system conduit. As such, two fluid flows may be directed out of the same first rocket engine vent system conduit in at least substantially opposite directions to at least substantially cancel any lateral thrust vectors that may be exerted on the first rocket engine vent system conduit. This also allows the fluid flows to be directed from the first rocket engine vent system conduit away from sensitive components or so as to otherwise reduce the potential for damaging other components of the rocket.
Axial thrusts to which the upper stage rocket engine is exposed during venting may be significantly reduced. In one embodiment, the end of the first rocket engine vent system conduit that is opposite that which interfaces with the upper stage rocket engine is closed. Stated another way, axial thrusts may be significantly reduced by having the projected area of the first and second vanes onto a plane perpendicular to a center reference axis about which the first rocket engine vent system conduit may be disposed, equal to the area of the first rocket engine vent system conduit defined by its diameter.
Multiple components may collectively define the first rocket engine vent system conduit associated with the subject first aspect. For instance, this first rocket engine vent system conduit may include a first conduit section having a first end that interfaces with the upper stage rocket engine. A separate first flow adapter may be separately interconnected with this first conduit section. The noted pair of vent apertures and the first flow diverter may be part of this first flow adapter. In one embodiment, the first flow adapter is attached to and extends beyond a second end of the first conduit section that is opposite the above-noted first end that interfaces with the upper stage rocket engine.
The first rocket engine vent system conduit may be disposed about a center reference axis and may include a pair of beveled surfaces. Each beveled surface may include one of the noted vent apertures. These beveled surfaces may be oriented on the first rocket engine vent system conduit so as to extend at least generally toward the noted center reference when progressing in a direction of a downstream end of the first rocket engine vent system conduit. In one embodiment, the pair of beveled surfaces are disposed in opposing relation, or stated another way, the beveled surfaces are spaced about 180 degrees apart. The beveled surfaces allow for increasing the surface area of the vent apertures for purposes of allowing adequate flow out of the first rocket engine vent system conduit.
The pair of vanes of the first flow diverter may direct two different fluid flows out of the same first rocket engine vent system conduit. Each fluid flow is directed through its own vent aperture (i.e., there is a one-to-one correspondence between the vanes of the first flow diverter and the vent apertures). Preferably, these fluid flows are at least substantially equal. Bifurcation of the fluid flow through the first rocket engine vent system conduit may be facilitated by configuring an upstream portion of the first flow diverter as an edge, and having the noted pair of vanes extend away from this edge at least generally away from a center reference axis about which the first rocket engine vent system conduit may be disposed. Configuring the upstream extreme of the first flow diverter as an edge also reduces the potential for the first flow diverter creating an undesirable back pressure within the first rocket engine vent system conduit.
Preferably the vanes extend at least generally away from a center reference axis about which the first rocket engine vent system conduit may be disposed to provide the desired flow diversion function. The vanes of the first flow diverter may be configured as arcuate surfaces to provide this flow diversion function. In one embodiment, each of the vanes of the first flow diverter is defined by a single radius. Other profiles may be appropriate for the vanes and which would still provide the desired flow diversion function and in a desired manner. For instance, the vanes may be an at least substantially planar surface (e.g., the first flow diverter could be wedge-shaped).
The first aspect may further include a second stage that is interconnected with the upper stage on an opposite end thereof in relation to the spacecraft, and further that includes at least one rocket engine. The upper stage rocket engine vent system may be configured so as to be able to discharge flows from the upper stage rocket engine even when the upper stage remains interconnected with the second stage, typically via an interstage adapter. In this regard, the first rocket engine vent system conduit may include first and second legs. The first leg may be structurally interconnected with the upper stage and may incorporate the noted pair of vent apertures and the first flow diverter. The second leg may be structurally interconnected with the second stage or an interstage adapter, and further may be disposed over the first leg when the second stage is interconnected with the upper stage. As such, any fluid flow that is directed out of the pair of vent apertures in the first leg is directed into the corresponding second leg when the second stage is interconnected with the upper stage. The second leg extends through an outer wall of the second stage or more typically an interstage adapter to direct any such fluid flow overboard. Preferably, the second leg is slidably interconnected with the first leg such that when the second stage separates from the upper stage, the second leg is simply xe2x80x9cpulled offxe2x80x9d its corresponding first leg such that the upper stage rocket engine vent system still may be used to discharge rocket fuel from the upper stage rocket engine. That is, the upper leg will remain structurally interconnected with the upper stage, and the xe2x80x9cpulling offxe2x80x9d of the second leg from the first leg will expose the noted pair of vent apertures such that rocket fuel may be discharged directly into outer space during flight.
A second aspect of the present invention is embodied by a rocket that includes a spacecraft, as well as an upper stage that is interconnected with this spacecraft and that includes at least one rocket engine. An upper stage rocket engine vent system is fluidly interconnected with this upper stage rocket engine and includes at least one rocket engine vent system conduit (hereafter a xe2x80x9cfirst rocket engine vent system conduitxe2x80x9d). This first rocket engine vent system conduit includes appropriate structure for directing a first fluid flow out of the first rocket engine vent conduit in a first direction that is disposed at an angle relative to a center reference axis about which the first rocket engine vent system conduit is disposed (e.g., at least generally transverse to such a center reference axis).
Various refinements exist of the features noted in relation to the subject second aspect of the present invention. Further features may also be incorporated in the subject second aspect of the present invention as well. These refinements and additional features may exist individually TS or in any combination. The first rocket engine vent system conduit may further include structure for directing a second fluid flow out of this conduit in a second direction that is also disposed at an angle relative to the center reference axis about which the first rocket engine vent system conduit is disposed, preferably simultaneously with the above-noted first fluid flow. More preferably, the first and second fluid flows are at least generally directly opposite each other to significantly reduce the lateral thrust to which the first rocket engine vent system is exposed during venting operations.
A third aspect of the present invention relates to a vent system for a high pressure system. The system includes a high pressure source and a vent system that is fluidly interconnected therewith. This vent system includes a first vent system conduit. A plurality of vent apertures extend entirely through this first vent system conduit. A first flow diverter is disposed within this first vent system conduit so as to direct a fluid flow that is progressing through the first vent system conduit in a radially outward direction (i.e., at least generally away from a center reference axis about which the first vent system conduit is disposed) toward the plurality of vent apertures when the first flow diverter is encountered by this fluid flow.
Various refinements exist of the features noted in relation to the subject third aspect of the present invention. Further features may also be incorporated in the subject third aspect of the present invention as well. These refinements and additional features may exist individually or in any combination. Initially, the various features of the subject third aspect may be used by both the first and second aspects discussed above. The static pressure profile throughout the first vent system conduit may be at least substantially unchanged (in relation to such a conduit that does not include the first flow diverter and vent apertures) by having the collective area of the plurality of vent apertures (i.e., the surface area occupied by these vent apertures) be greater than the cross-sectional area of the first vent system conduit (i.e., the area of the first vent system conduit, taken perpendicular to a center reference axis about which the first vent system conduit may be disposed). If a significant back pressure did develop in the first vent system conduit, that is if the venting did not occur quickly enough such that this conduit (as well as the high pressure source fluidly interconnected therewith) became overpressurized or such that the static pressure therein significantly increased, one or more components of the high pressure system could become damaged. Reduction of the lateral thrust to which the first vent system conduit may be exposed during a venting of the high pressure source may be realized by disposing the plurality of vent apertures about the first vent system conduit in at least substantially equally spaced relation, or by disposing the plurality of vent apertures in symmetrical fashion on the first vent system conduit.
Multiple and discrete surfaces may be utilized by the first flow diverter for diverting the fluid flow to the plurality of vent apertures. Each discrete surface may be associated with its own vent aperture, or a single discrete surface of the first flow diverter may be associated with a plurality of vent apertures. The first flow diverter may also be in the form of a single, continuous surface. For instance, the first flow diverter may be in the form of a cone or the like, with the vertex of the cone being disposed so as to project upstream.
Diversion of the fluid flow toward the plurality of vent apertures may be affected by configuring the first flow diverter so as to extend at least generally away from a center reference axis about which the first vent system conduit may be disposed in a direction of the flow through the first vent system conduit. The first flow diverter may also actually separate the fluid flow into a plurality of separate and discrete fluid flows.
A fourth aspect of the present invention is embodied by a multiple stage rocket. Each stage of the rocket includes at least one rocket engine. A first vent system is fluidly interconnected with a high pressure source (e.g., rocket engine, oxidizer system, propellant tank) that is associated with a first stage of the rocket. This first vent system includes a first leg that is fluidly interconnected with this high pressure source, and further that is physically interconnected with the first stage. Multiple vent apertures are provided on the first leg so as to direct a plurality of fluid flows in an at least generally radially outward direction relative to a center axis about which the first leg is disposed. A second leg of the vent system is fluidly interconnected with the first leg, and is physically interconnected with a second stage of the rocket or some interconnecting structure between the first and second stages (e.g., an interstage adapter). While the second stage is interconnected with the first stage, venting of the high pressure source is directed out of the second leg. The resultant fluid flow from this venting may flow through the first leg and then into the second leg. After the second stage has separated from the first stage and as the first stage continues to progress through outer space, venting of the high pressure source is directed out of the first leg and the noted plurality of vent apertures directly into outer space. Those features discussed above in relation to the first aspect of the present invention may be utilized by this fourth aspect of the invention and in any combination.